National Advisory Committee for Aeronautics
N0.91
R
esearc
h Abstracts
OCTOBER 21, 1955
CURRENT NACA REPORTS
NACA Rept. 1191
ON THE DEVELOPMENT OF TURBULENT WAKES
FROM VORTEX STREETS. Anatol Roshko, Cali-
fornia Institute of Technology. 1954. ii, 25p.
diagrs., photos., 3 tabs. (NACA Rept. 1191.
Formerly TN 2913)
Wake development behind circular cylinders at
Reynolds numbers from 40 to 10,000 was investigated
by hot-wire techniques in a low-speed wind tunnel.
The Reynolds, number range of periodic vortex
shedding is divided into two distinct subranges. In
the stable range, R = 40 to 150, regular vortex
streets are formed and no turbulent motion develops,
the vortices decaying by viscous diffusion. The
range.R = 150 to 300 is a transition region to the
irregular range in which turbulent velocity fluctua-
tions accompany the periodic formation of vortices.
The diffusion is turbulent and the wake becomes
fully turbulent in 40 to 50 diameters. The turbulence
is initiated by laminar-turbulent transition in the
free layers which spring from the separation points
on the cylinder. An annular vortex street was ob-
served in the wake of a ring.
NACA RM E55E18
PERFORMANCE CHARACTERISTICS OF HEMI-
SPHERICAL TARGET-TYPE THRUST REVERSERS.
Fred W. Steffen, Jack G. McArdle, and James W.
Coats. September 1955. 39p. diagrs., photos.,
tab. (NACA RM E55E18)
Reverse-thrust performance of hemispherical
target -type thrust reversers was obtained. The
value of reverse-thrust ratio was found to be pri-
marily a function of hemisphere diameter. Slight
improvements in performance were obtained with
some size reversers by operating with the exhaust
nozzle opened slightly. When high values of
reverse-thrust ratio were obtained, the reversed
flow attached to the boattail and the pressures on the
boattail fell belovv atmospheric. This pressure re-
duction amounted to about 20 percent of the resultant
reverse force. Various boattail shapes had little ef-
fect on reverse-thrust ratio for most hemisphere
sizes.
NACA RM E55G27a
IDEAL TEMPERATURE RISE DUE TO CONSTANT-
PRESSURE COMBUSTION OF A JP-4 FUEL.
S(idney) C. Huntley. September 1955. 53p.
diagrs., tabs. (NACA RM E55G27a)
Charts are presented from which ideal temperature
rise or the ideal quantity of fuel required to obtain a
specified combustion temperature may be obtained.
The charts are applicable only to a fuel having a
hydrogen-carbon mass ratio of 0.168 (CH2) and in-
clude a range of fuel-air ratios from to 1.2 frac-
tions of stoichiometric fuel-air ratio >vith dissocia-
tion taken into account, inlet-air temperatures from
400° to 1600° R, and combustion pressures from'
1/16 to 64 atmospheres
NACA RM L55G21
NOTE ON HOVERING TURNS WITH TANDEM
HELICOPTERS. John P. Reeder and Robert J.
Tapscott. September 1955. 5p. photo. (NACA
RM L55G21)
The source of an appreciable pitching-moment dif-
ference between left and right hovering turns for a
tandem helicopter is described. The difference in
pitching moment results from the difference in ro-
tational speed of the counterrotating rotors with
respect to the air while the helicopter is turning.
NACA TM 1377
THE THEORIES OF TURBULENCE. (Les Theories
de la Turbulence). L. Agostini and J. Bass.
October 1955. 163p. diagrs. (NACA TM 1377.
Trans, from Ministere de l'Air, Publications
Scientifiques et Techniques 237, 1950)
The report includes a discussion of the kinematics
of statistical mediums, particularly those which are
isotropic. A mathematical study is made of the ap-
plications of Navier's equations to turbulent motion.
Physical theories involving similarity are dealt with.
Review is made of much of the work in turbulence.
The theoretical discussions are illustrated by some
correlation and spectrum curves based on measure-
ments taken in the wind tunnel at the laboratory of
the mechanics of the atmosphere at Marseille.
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NACA TM 1389
OPTIMUM FLIGHT PATHS OF TURBOJET AIR-
CRAFT. (Traiettorie Ottime Di Volo Degli Aero-
plani Azionati Da Turboreattori). Angelo Miele.
September 1955. 47p. diagrs. , tabs. (NACA TM
1389. Trans, from L'Aerotecnica, v. 32, no. 4,
1952, p. 206-219)
The climb of turbojet aircraft is analyzed and dis-
cussed including the effects of tangential accelera-
tions. Three flight performances are examined:
minimum time of climb, climb with minimum fuel
consumption, and steepest climb. Diagrams for
quick calculation of the optimum Mach numbers and
the effect of acceleration on the rate of climb in
tropospheric and stratospheric flight are given.
NACA TM 1396
FROM LINEAR MECHANICS TO NONLINEAR
MECHANICS. (De la mecanique lineaire a la
mecanique non lineaire). Julien Loeb. October
1955. 18p. diagrs.^ ^NACA TM 1396. Trans,
from Annales des Telecommunications, v. 5, no. 2,
Feb., 1950, p. 65-71)
Consideration is first given to the technique used in
telecommunication where a nonlinear system (the
modulator) results in a linear transposition of a
signal. It is then shown that a similar method per-
mits linearization of electromechanical devices or
nonlinear mechanical devices. A sweep function
plays the same role as the carrier wave in radio-
electricity. The linearizations of certain non-
linear functionals are presented.
NACA TN 3463
INVESTIGATION OF THE VIBRATIONS OF A HOL-
LOW THIN-WALLED RECTANGULAR BEAM.
Eldon E. Kordes and Edwin T. Kruszewski. October
1955. 24p. diagrs., photos., 2 tabs. (NACA
TN 3463)
Experimental modes and frequencies of an unstif-
fened hollow beam of rectangular cross section are
presented, and comparisons are made between ex-
perimental and theoretical frequencies. Theories
based on rigid cross sections were found to be suf-
ficiently accurate to predict the frequencies of only
the lovvest three bending modes. For the higher
bending modes and all the torsional modes, it was
necessary to include the effects of cross-sectional
distortions in the calculations.
NACA TN 3464
INFLUENCE OF SHEAR DEFORMATION ON THE
CROSS SECTION ON TORSIONAL FREQUENCIES
OF BOX BEAMS. Edwin T. Kruszewski and William
W. Davenport. October 1955. 23p. diagrs. (NACA
TN 3464)
An exact analysis has been carried out on the tor-
sional vibrations of a four-flange box beam with
NACA
RESEARCH ABSTRACTS NO. 91
cross sections which can change shape because the
stiffness of the bulkheads is finite. The effect of
shear deformation of the cross section on the tor-
sional frequencies is illustrated by numerical cal-
culations. An approximate method for quickly
estimating the effects of bulkhead shear stiffness on
the torsional frequencies of box beams has been
devised.
NACA TN 3475
AN ANALYSIS OF ACCELERATION, AIRSPEED,
AND GUST-VELOCITY DATA FROM ONE TYPE OF
FOUR-ENGINE TRANSPORT AIRPLANE OPERATED
OVER TWO DOMESTIC ROUTES. Martin R. Copp
and Thomas L. Coleman. October 1955. 29p.
diagrs., tabs. (NACA TN 3475)
Time-history data obtained by the NACA VGH re-
corder from one type of four-engine commercial
transport airplane during operations on two domestic
routes indicated that the number of gust accelera-
tions experienced per mile of flight by the two opera-
tions differed by a factor of roughly 3. The number
of gusts per mile of flight differed by a factor of
roughly 4. A general decrease in the frequency of
occurrence of gust velocities with increasing altitude
was noted for both operations. For acceleration
values above 0.8g, maneuver accelerations formed a
substantial part of the total-flight load histories. A
comparison of the average overall airspeeds and the
corresponding average airspeeds in rough air (with
accelerations equal to or greater than 0.3g) indicated
very little slowdown by either operation upon en-
countering turbulence.
NACA TN 3481
WIND-TUNNEL INVESTIGATION AT LOW SPEED
OF EFFECT OF SIZE AND POSITION OF CLOSED
AIR DUCTS ON STATIC LONGITUDINAL AND
STATIC LATERAL STABILITY CHARACTERISTICS
OF UNSWEPT -MID WING MODELS HAVING WINGS
OF ASPECT RATIO 2, 4, AND 6. Byron M. Jaquet
and James L. Williams. September 1955. 45p.
diagrs., photos., tabs. (NACA TN 3481)
Results are presented of tests made at a Mach num-
ber of 0.13 in the Langley stability tunnel to deter-
mine the effects of closed wing-root air ducts on the
static longitudinal and static lateral stability charac-
teristics of models having unswept wing and tail sur-
faces with wings of aspect ratio 2, 4, and 6. In
addition, for model configurations employing the
wing of aspect ratio 2 the effects of top and bottom
fuselage ducts on the static longitudinal and static
lateral characteristics were determined. The effect
of the wing-root ducts on the aerodynamic hystere-
sis in sideslip of the model employing the wing of
aspect ratio 2 was also determined.
NACA
RESEARCH ABSTRACTS NO. 91
NACA TN 3485
AN APPROXIMATE SOLUTION FOR AXIALLY
SYMMETRIC FLOW OVER A CONE WITH AN AT-
TACHED SHOCK WAVE. Richard A. Hord.
October 1955. 32p. diagrs. (NACA TN 3485)
It is shown that the streamlines in an angular
neighborhood of the surface of an unyawed circular
cone with an attached shock wave are, to a first ap-
proximation, portions of hyperbolas. This fact is
used as a basis for the development of an approxi-
mate solution in which shock-wave orientation and
flow field behind the shock wave are given explicitly
in terms of free-stream Mach number, vertex angle
of the body cone, and the ratio of specific heats of
the gas. The approximate solution is compared
with other approximate solutions for the cone.
NACA TN 3497
SUMMARY OF RESULTS OF A WIND-TUNNEL
INVESTIGATION OF NINE RELATED HORIZONTAL
TAILS. Jules B. Dods, Jr. and Bruce E. Tinling.
July 1955. 105p. diagrs., 2 tabs. (NACA TN 3497.
Formerly RM A51G31a)
A compilation of data is presented for models of
nine related horizontal tails. The majority of the
results were obtained at a Mach number of approxi-
mately 0.20. Three of the models were tested
throughout the subsonic Mach number range to a
maximum of 0.94. The Reynolds number range was
from 2 to 4 million. The models had aspect ratios
from 2 to 6, angles of sweepback from 5. 7° to 45°,
and had 30-percent-chord, sealed, plain flaps. The
lift coefficient, hinge-moment coefficient, and pres-
sure coefficients across the elevator nose seal are
presented. The effects of sweepback, aspect ratio,
and Mach number on the lift and hinge-moment
parameters are summarized. Comparisons of the
experimental results with theoretical calculations
are presented.
NACA TN 3519
VISUALIZATION STUDY OF SECONDARY FLOWS
IN TURBINE ROTOR TIP REGIONS. Hubert W.
Allen and Milton G. Kofskey. September 1955.
33p. diagrs., photos., tab. (NACA TN 3519)
A low-speed visualization study of turbine rotor tip
secondary flows was made. Results include qualita-
tive information on tip-clearance flow, cross-
passage flow, and scraping flow, and on a range of
rotor speeds for which a transition condition appear-
ed with minimum flow disturbance. Rotor speed re-
quired for this transition flow condition depended
upon blade configuration, angle of incidence, tip
clearance, and air-flow rate. Results should aid in
extending the study to higher airspeeds.
NACA TN 3532
LOW-SPEED STATIC LATERAL AND ROLLING
STABILITY CHARACTERISTICS OF A SERIES OF
CONFIGURATIONS COMPOSED OF INTERSECTING
TRIANGULAR PLAN-FORM SURFACES. David F.
Thomas, Jr. October 1955. 29p. diagrs., photos.
(NACA TN 3532)
The static lateral and rolling stability derivatives
of a series of cruciform, inverted T-, V-, and Y-
configurations composed of low-aspect-ratio trian-
gular surfaces have been obtained at low speed in
the 6-foot-diameter rolling-flow test section of the
Langley stability tunnel. These derivatives are
presented as functions of the geometry of the models,
and for two configurations (a planar wing and an in-
verted T), as functions of angle of attack. Where
possible, comparisons have been made to indicate
the extent of agreement between experiment and
existing theory. In general, the sideslip deriva-
tives showed better agreement between theory and
experiment than the rolling derivatives.
NACA TN 3533
THE PROPER COMBINATION OF LIFT LOADING
FOR LEAST DRAG ON A SUPERSONIC WING.
Frederick C. Grant. October 1955. 21p. diagrs.
(NACA TN 3533)
Lagrange's method of undetermined multipliers is
applied to the problem of properly combining lift
loadings for the least drag at a given lift on super-
sonic wings. The interference drag between the
optimum loading and any loading at the same lift
coefficient is found to be constant on a given plan
form. This is an integral form of a criterion
established by Robert T. Jones for optimum load-
ings. The best combination of four loadings on a
delta wing with subsonic leading edges is calculated
for several Mach numbers as a numerical example.
NACA TN 3535
FLIGHT INVESTIGATION OF THE SURFACE-
PRESSURE DISTRIBUTION AND THE FLOW FIELD
AROUND A CONICAL AND TWO SPHERICAL NON-
ROTATING FULL-SCALE PROPELLER SPINNERS.
Jerome B. Hammack, Milton L. Windier, and
Elwood F. Scheithauer. September 1955. 36p.
diagrs. , photos. (NACA TN 3535)
The surface-pressure distribution and the flow field
around a conical and two spherical nonrotating full-
scale propeller spinners were determined in flight
at Mach numbers of 0. 70 to 0.96. The local-surface
Mach numbers between the cone-sphere tangency and
the maximum thickness stations were approximately
0.4 higher than those at a free-stream Mach number
of 0.95. The departure from free-stream conditions
of the larger spherical spinner extends beyond the
1. 3-spinner-radius station; whereas, with the
smaller spherical spinner, free-stream conditions
were reached at the 1. 3-spinner-radius station.
NACA TN 3536
A LIMITED FLIGHT INVESTIGATION OF THE EF-
FECT OF THREE VORTEX-GENERATOR CONFIGU-
RATIONS ON THE EFFECTIVENESS OF A PLAIN
FLAP ON AN UNSWEPT WING. Garland J. Morris
and Lindsay J(ohn) Lina. September 1955. 20p.
diagrs. , photos. , tabs. (NACA TN 3536)
An exploratory flight investigation was made to deter-
mine the effect of three vortex-generator configura-
tions on the effectiveness of the plain flap of a fighter
airplane. Tests were made with flaps deflected 19°
and 45° at several indicated airspeeds in the range
from stall to 140 miles per hour. No improvement
in flap effectiveness was obtained with the flaps de-
flected full down. With flaps deflected 19°, the
vortex-generator configuration which produced the
largest increase in lift coefficient had about the same
effect as increasing the flap deflection to about 27°.
Another configuration produced some increase in
lift coefficient with no apparent increase in drag.
NACA TN 3538
SUMMARY OF DERIVED GUST VELOCITIES OB-
TAINED FROM MEASUREMENTS WITHIN THUN-
DERSTORMS. H(arold) B. Tolefson. October
1955. 19p. diagrs., tabs. (NACA TN 3538)
Available measurements of the derived gust veloci-
ties within thunderstorms are summarized for alti-
tudes from 5,000 to 34,000 feet. The results indi-
cate that the intensity of the derived gust velocity is
essentially constant up to altitudes of 20,000 feet and
that an approximate 10-percent reduction in the gust
intensity occurs for altitudes from 20,000 to 30,000
feet.
NACA TN 3561
INTENSITY, SCALE, AND SPECTRA OF TURBU-
LENCE IN MIXING REGION OF FREE SUBSONIC
JET. James C. Laurence. September 1955. 58p.
diagrs., photo., tab. (NACA TN 3561)
Hot-wire anemometer measurements of the turbu-
lence parameters were made in a 3.5-inch-diameter
free jet at exit Mach numbers between 0.2 and 0.7
and Reynolds numbers (based on jet rad. ) between
37,500 and 350,000. The results of these measure-
ments show that (1) the intensity of turbulence is a
max. at a distance of approximately 1 jet rad. from
the jet center line and decreases with increasing
Mach and/or Reynolds number, and (2) the lateral
and longitudinal scales of turbulence are nearly in-
dependent of Mach and/or Reynolds number and vary
proportionally with distance from the jet nozzle.
The lateral scale is much smaller than the longi-
tudinal and does not vary with distance from the
center line, while the longitudinal scale is a max.
at a distance from the center line of about 0.7 to 0.8
of the jet rad.
NACA
RESEARCH ABSTRACTS NO. 91
NACA TN 3568
AVERAGING OF PERIODIC PRESSURE PULSA-
TIONS BY A TOTAL-PRESSURE PROBE. R. C.
Johnson. October 1955. 30p. diagrs., photo.,
tabs. (NACA TN 3568)
Information is presented on the average pressure
indicated by a total -pressure probe subjected to a
stagnation pressure that alternates periodically be-
tween two constant values. Calculated and experi-
mental data are in good agreement, and errors are
reduced when the probe design is such as to ensure
laminar-flow pulsations in the probe at all times.
The averaging error is minimized when the inside
diameter of the probe entrance tube is made as
small as possible, and its length as great as possi-
ble, consistent with an acceptable time lag.
NACA TN 3570
AN EXPERIMENTAL COMPARISON OF THE
LAGRANGIAN AND EULERIAN CORRELATION
COEFFICIENTS IN HOMOGENEOUS ISOTROPIC
TURBULENCE. William R. Mickelsen. October
1955. 42p. diagrs. (NACA TN 3570)
The Lagrangian and Eulerian correlation coeffi-
cients were compared in a field of homogeneous,
isotropic turbulence. The Lagrangian correlation
coefficient was characterized by diffusion measure-
ments, and the Eulerian coefficient was measured
by hot-wire anemometry. The Lagrangian and
Eulerian correlation coefficients had similar shapes
connected by a linear relation between their coordi-
nates. The proportionality factor in the linear rela-
tion was roughly constant over a range of turbulence
intensities from 1.8 to 14 feet per second. The lin-
ear relation permits solution of mixing problems
from the Eulerian turbulence parameters.
BRITISH REPORTS
N-38869*
Royal Aircraft Establishment (Gt. Brit.)
A SIMPLIFIED MODEL OF THE INCOMPRESSIBLE
FLOW PAST TWO-DIMENSIONAL AEROFOILS
WITH A LONG BUBBLE TYPE OF FLOW SEPARA-
TION. J. F. Norbury and L. F. Crabtree.
June 19,55. 17p. diagrs., tabs. (RAE Tech. Note
Aero 2352)
This problem may be divided into two parts. The
first part concerns the external inviscid flow, and
the other is related to the details of the flow inside
the viscous region formed by the bubble, the bound-
ary layer, and the wake. Although much informa-
tion can be obtained by considering only the external
stream (as in the hodograph method developed by
Maskell), a complete and unique solution for the
pressure distribution can only be found by attacking
the second part of the problem as well. This has
been attempted here and it is shown how the pres-
sure recovery ratio at the end of a long bubble may
be found by the analysis of a simplified model of the
flow; the complete pressure distribution may then
be uniquely determined.
NACA
RESEARCH ABSTRACTS NO. 91
MISCELLANEOUS
N-39013*
AIRCRAFT STRUCTURES RESEARCH AT ELEVAT-
ED TEMPERATURES. John E. Duberg. (Present-
ed to Structures and Materials Panel of Advisory
Group for Aeronautical Research and Development
(NATO) London, England, September 5-9, 1955).
36p. diagrs. , photos.
A review is made of the test techniques that have
been developed and used by the NACA for experimen-
tal research in aircraft structures at elevated temp-
eratures. Some experimental results are presented.
Remarks are included on the problem of model scal-
ing for testing of structures at high temperatures.
UNPUBLISHED PAPERS
N-37996*
THE BUCKLING OF PLATES AND BARS IN THE
PLASTIC RANGE. I. THEORY. (Over het
knitvraagstuk in het plastische gebied bij staven en
platen. I. Theorie). J. F. Besseling. July 1955.
97p. diagrs., tabs. (Trans, from Nationaal
Luchtvaartlaboratorium, Amsterdam, S.407, Oct. 14,
1952)
The topics treated include the plastic buckling of
bars and plates; application of plasticity theories to
'the determination of stiffness quantities needed for
computing the buckling; and the effect of plasticity
on the buckling load for a number of fundamental
buckling cases when the buckling stress condition is
reached through a uniform increase of loading. This
report is limited to the theoretical side of the buck-
ling problem and gives the basis for comparing
theoretical and experimental results.
DECLASSIFIED NACA REPORTS
NACA RM A8I17
INVESTIGATION OF A THIN WING OF ASPECT
RATIO 4 IN THE AMES 12-FOOT PRESSURE WIND
TUNNEL. Ill - THE EFFECTIVENESS OF A
CONSTANT-CHORD AILERON. Ben H. Johnson, Jr.
and Fred A. Demele. November 19, 1948. 26p.
diagrs., photo. (NACA RM A8I17)
(Declassified from Confidential, 9/15/55)
Presented in the report are results of tests at Mach
numbers from 0.27 to 0.94 of a thin, unswept wing
having a modified diamond airfoil section of thickness
ratio 0.042 and equipped with a constant-chord ailer-
on. The tests were conducted at a constant Reynolds
number of 2, 730, 000. The effects of compressibility
on the aileron effectiveness were negligible at Mach
numbers up to 0.85, but at higher Mach numbers er-
ratic effects of compressibility were evident, espe-
cially at lift coefficients greater than 0.5.
NACA RM A9I01
INVESTIGATION OF A THIN WING OF ASPECT
RATIO 4 IN THE AMES 12-FOOT PRESSURE WIND
TUNNEL. V - STATIC LONGITUDINAL STABILITY
AND CONTROL THROUGHOUT THE SUBSONIC
SPEED RANGE OF A SEMISPAN MODEL OF A
SUPERSONIC AIRPLANE. Ben H. Johnson, Jr. .
and Francis W. Rollins. December 8, 1949. 130p.
diagrs., photos. (NACA RM A9I01)
(Declassified from Confidential, 9/15/55)
Presented in this report are results of tests at Mach
numbers from 0.20 to 0.94 of a model of a hypo-
thetical supersonic airplane equipped with a thin,
sharp-edged wing and tail without sweep. The static
longitudinal-stability characteristics of the model
have been measured for two different vertical loca-
tions of the horizontal tail. The longitudinal control
afforded by an all-movable stabilizer and by an ele-
vator has been investigated. The downwash at the
tail has been computed and the dynamic-pressure
ratio at the tail has been evaluated from pressure
measurements in the wake of the wing.
NACA RM A51I07
THE STATIC LONGITUDINAL CHARACTERISTICS
AT MACH NUMBERS UP TO 0.95 OF A
TRIANGULAR-WING CANARD MODEL HAVING A
TRIANGULAR CONTROL. Jack D. Stephenson and
Ralph Selan. December 1951. 72p. diagrs., photo.
(NACA RM A51I07) (Declassified from
Confidential, 9/15/55)
Presents and analyzes results of tests to assess the
longitudinal characteristics of a canard-type model
having a triangular wing of aspect ratio 2 and NACA
0008-63 sections. The horizontal canard surface
had a plan form identical to the wing and an NACA
0005-63 section. The Mach number range of the
investigation varied from 0.25 to 0.95 at Reynolds
numbers of 8 million and 3 million, respectively.
The model was tested with the horizontal control
surface at various fixed angles of incidence and with
the surface unrestrained so that it could pivot about
an axis at 30 percent of its mean aerodynamic chord.
NACA RM A52F18
THE LONGITUDINAL CHARACTERISTICS AT MACH
NUMBERS UP TO 0.92 OF A CAMBERED AND
TWISTED WING HAVING 40° OF SWEEPBACK AND
AN ASPECT RATIO OF 10. George G. Edwards,
Bruce E. Tinling, and Arthur C. Ackerman.
September 1952. 71p. diagrs., photos., tab.
(NACA RM A52F18) (Declassified from
Confidential, 9/15/55)
A sweptback wing, in combination with a fuselage, of
a type considered suitable for long-range, high-
speed airplanes, has been tested in the Ames 12-foot
pressure wind tunnel. The wing had 40° of sweep-
back, an aspect ratio of 10, a taper ratio of 0.4, and
5° of washout at the tip. The sections normal to the
sweep reference line had NACA four-digit profiles,
design lift coefficients of 0.40, and varied in thick-
ness ratio from 14 percent at the root to 11 percent
at the tip. The lift, drag, and pitching moment of a
semispan model were measured at Reynolds num-
bers from 2,000,000 to 8,000.000 at low Mach num-
bers, and at Mach numbers from 0.25 to 0.92 at a
Reynolds number of 2,000,000. The changes in
boundary-layer flow on the upper surface were
studied with tufts.
NACA RM A52I19
THE LONGITUDINAL CHARACTERISTICS AT MACH
NUMBERS UP TO 0.9 OF A WING-FUSELAGE-TAIL
COMBINATION HAVING A WING WITH 40° OF
SWEEPBACK AND AN ASPECT RATIO OF 10.
Bruce E. Tinling. December 1952. 41p. diagrs.,
photo., tab. (NACA RM A52I19) (Declassified from
Confidential, 9/15/55)
Wind-tunnel tests were made of a wing-fuselage hori-
zontal tail combination suitable for long-range high-
speed airplanes. The cambered and twisted wing
had an aspect ratio of 10, taper ratio of 0.4, and 40°
sweepback. The all-movable horizontal tail had an
aspect ratio of 4.5, taper ratio of 0.4, and 40°
sweepback. Included are data on longitudinal sta-
bility and control, wing fence and tail effect, and
isolated horizontal tail data. Data »vere obtained at
a Reynolds number of 8,000,000 at low Mach num-
bers, and at a Reynolds number of 2,000,000 at Mach
numbers up to 0.9.
NACA RM A52K20
PRESSURE DISTRIBUTION AT MACH NUMBERS UP
TO 0.90 ON A CAMBERED AND TWISTED WING
HAVING 40° OF SWEEPBACK AND AN ASPECT
RATIO OF 10, INCLUDING THE EFFECTS OF
FENCES. Frederick W. Boltz and Harry H.
Shibata. March 1953. 133p. diagrs., photos., tabs.
(NACA RM A52K20) (Declassified from
Confidential, 9/15/55)
Pressure-distribution measurements have been made
on a semispan model of a cambered and twisted wing,
NACA
RESEARCH
ABSTRACTS NO. 9)
alone and in combination with a fuselage. The wing
had 40° of sweepback, an aspect ratio of 10, a taper
ratio of 0.4, and 5° of washout at the tip. The wing
thickness distribution in sections normal to the ref-
erence sweep line was the NACA 4-digit series with
the maximum thickness varying from 14-percent
chord at the root to 11-percent chord at the tip. The
chordwise distributions of pressure coefficient at
nine semispan stations are presented for Mach num-
bers of 0.165 and 0.25 at a Reynolds number of
8,000,000, and for Mach numbers from 0.25 to 0.90
at a Reynolds number of 2,000,000. Tabulated
pressure data are presented for the wing without
fences and with four fences.
NACA RM A53D06
THE EFFECTS OF NACELLES AND OF EXTENDED
SPLIT FLAPS ON THE LONGITUDINAL CHARAC-
TERISTICS OF A WING-FUSELAGE-TAIL COMBI-
NATION HAVING A WING WITH 40° OF SWEEP-
BACK AND AN ASPECT RATIO OF 10. Bruce E.
Tinling and Armando E. Lopez. June 1953. 47p.
diagrs., tab. (NACA RM A53D06) (Declassified
from Confidential, 9/15/55)
Wind-tunnel tests were made to evaluate the effects
of nacelles and of extended split flaps on the longitu-
dinal characteristics of a wing-fuselage-tail combi-
nation suitable for long-range high-speed airplanes.
The cambered and twisted wing had an aspect ratio
of 10, a taper ratio of 0.4, and 40° of sweepback.
The nacelles were at 25 and 50 percent of the semi-
span. Data were obtained to evaluate the effects of
nacelles at Mach numbers up to 0.90 at a Reynolds
number of 2,000,000. The effects of the flaps were
evaluated from data obtained at a Reynolds number
of 4,000,000 and a Mach number of 0.082. A
limited number of data were also obtained to evalu-
ate the effects of increasing the tail height.
NACA RM A53I23
DOWNWASH BEHIND A TRIANGULAR WING OF
ASPECT RATIO 3 - TRANSONIC BUMP METHOD.
John A. Axelson. December 1953. 37p. diagrs.,
photo., tab. (NACA RM A53I23) (Declassified from
Confidential, 9/15/55)
Downwash measured by means of an all-movable
horizontal tail in several different locations is pre-
sented for a triangular wing having an aspect ratio of
3 and the NACA 63A006 section. The lift, drag, and
pitching-moment characteristics of the wing and the
downwash are presented for angles of attack up to
28° over a Mach number range from 0.6 to 1.1,
corresponding to a Reynolds number range from 1.8
million to 2.4 million. The effects of tail location
on the tail contribution to the static longitudinal sta-
bility of the wing-tail combination are discussed.
NACA
RESEARCH ABSTRACTS NO. 91
NACA RM A53I28
THE RESULTS OF WIND-TUNNEL TESTS AT LOW
SPEEDS OF A FOUR-ENGINE PROPELLER-DRIVEN
AIRPLANE CONFIGURATION HAVING A WING WITH
40° OF SWEEPBACK AND AN ASPECT RATIO OF
10. George G. Edwards, Donald A. Buell, and
Jerald K. Dickson. December 1953. 121p. diagrs.,
photo., tabs. (NACA RM A53I28) (Declassified
from Confidential, 9/15/55)
The effects of operating propellers on the low-speed
longitudinal characteristics of a four-engine tractor
airplane configuration having a sweptback wing have
been investigated in the Ames 12-foot pressure wind
tunnel at thrust coefficients up to 0.9 per propeller
and at Reynolds numbers from 4,000,000 to
8,000,000. Variations in the model included different
heights and incidences of the horizontal tail as well
as tail removed, two arrangements of extended split
flaps, several propeller-blade angles, and inde-
pendent as well as simultaneous operation of the in-
board and outboard propellers. Coefficients of lift,
longitudinal force, pitching moment, propeller
thrust, and propeller power are presented in tabular
form for various values of advance ratio at constant
angles of attack. Selected portions of the data are
presented in plotted form for various constant thrust
coefficients.
NACA RM A53I29
THE TRANSONIC CHARACTERISTICS OF 36 SYM-
METRICAL WINGS OF VARYING TAPER, ASPECT
RATIO, AND THICKNESS AS DETERMINED BY THE
TRANSONIC -BUMP TECHNIQUE. Warren H.
Nelson, Edwin C. Allen, and Walter J. Krumm.
December 1953. 131p. diagrs., photo. (NACA
RM A53I29) (Declassified from Confidential,
9/15/55)
An investigation was made in the Ames 16-foot high-
speed wind tunnel, utilizing the transonic-bump tech-
nique to determine the effects of plan-form taper on
a series of wings having aspect ratios of 4, 3, and 2,
and NACA 63A00X sections with thickness-to-chord
ratios of 8, 6, 4, and 2 percent. The Mach number
range was 0.6 to 1.1 with a corresponding Reynolds
number range of about 1.4 million to 2.0 million.
The results indicate that increasing the taper ratio
caused only small increases in lift-curve slope
except for the wings of highest aspect ratio and
thinnest sections. Increasing taper ratio generally
increased the overall center-of-pressure travel in
going from subsonic to supersonic speeds.
NACA RM E8I01
EFFECT OF THREE FLAME -HOLDER CONFIGU-
RATIONS ON SUBSONIC FLIGHT PERFORMANCE
OF RECTANGULAR RAM JET OVER RANGE OF
ALTITUDES. Dugald O. Black and Wesley E.
Messing. November 24, 1948. 28p. diagrs., photo.,
tab. (NACA RM E8I01)
(Declassified from Confidential, 9/15/55)
A flight investigation is reported that was conducted
on a rectangular ram jet incorporating various
flame-holder configurations over a range of fuel -air
ratios from 0.017 to 0.120, combustion-chamber-
inlet velocities from 50 to 125 feet per second, and
pressure altitudes from 1500 to 28,000 feet. Highest
combustion efficiencies, which varied from a maxi-
mum of 82 percent at 1500 feet to 39 percent at
26,000 feet, were obtained at all altitudes with the
three-V gutter-type flame holder. However, at any
given altitude and fuel-air ratio slightly higher net-
thrust coefficients occurred with the two-V flame
holder as a result of its lower value of pressure
drop.
NACA RM E8I28
EFFECT OF VARIATION IN FUEL PRESSURE ON
COMBUSTION PERFORMANCE OF RECTANGULAR
RAM JET. Wesley E. Messing and Dugald O. Black.
November 24, 1948. 26p. diagrs., photo., tab.
(NACA RM E8I28)
(Declassified from Confidential, 9/15/55)
Reports effect of variation in fuel pressure on start-
ing characteristics, minimum blow-out limits, com-
bustion efficiencies, gas total-temperature ratio, and
net -thrust coefficient of rectangular ram jet operated
over range of pressure altitudes from 1500 to 26, 300
feet, indicated airspeeds from 100 to 200 miles per
hour, and fuel-air ratios from 0.017 to 0.120. In
general, increasing the degree of fuel atomization
and distribution by utilization of small orifice fuel
nozzles that operated at high fuel pressures resulted
in higher values of combustion efficiency, gas total-
temperature ratio, and net-thrust coefficient at a
given fuel -air ratio.
NACA RM E50D28
AERODYNAMIC CHARACTERISTICS OF NACA RM-
10 MISSILE IN 8- BY 6-FOOT SUPERSONIC WIND
TUNNEL AT MACH NUMBERS FROM 1.49 TO 1.98.
H - PRESENTATION AND ANALYSIS OF FORCE
MEASUREMENTS. Fred T. Esenwein, Leonard J.
Obery, and Carl F. Schueller. July 21, 1950. 34p.
diagrs., photo. (NACA RM E50D28)
(Declassified from Confidential, 9/15/55)
Experimental investigation of aerodynamic forces
acting on body of revolution (NACA RM-10 missile)
with and without stabilizing fins was conducted at
Mach numbers from 1.49 to 1.98 at angles of attack
from 0° to 9° and at Reynolds number of approxi-
mately 30, 000, 000. Comparison of experimental
lift, drag, and pitching-moment coefficients and
center-of-pressure location for body alone is made
with linearized potential theory and a semiempirical
method. Results indicate that aerodynamic charac-
teristics were predicted more accurately by semi-
empirical method than by potential theory. Break-
down of measured drag coefficients into components
of friction, pressure, and base-pressure drag is
presented for body alone at zero angle of attack.
NACA RM E 5011 9
AERODYNAMIC CHARACTERISTICS OF NACA
RM-10 MISSILE IN 8- BY 6-FOOT SUPERSONIC
WIND TUNNEL AT MACH NUMBERS FROM 1.49 TO
1.98. m - ANALYSIS OF FORCE DISTRIBUTION
AT ANGLE OF ATTACK (STABILIZING FINS RE-
MOVED). Roger W. Luidens and Paul C. Simon.
December 12, 1950. 26p. diagrs.
(NACA RM E50I19)
(Declassified from Confidential 9/15/55)
Analysis of force distribution on slender pointed
body of revolution at angle of attack was made utiliz-
ing pressure-distribution data and balance measure-
ments obtained in NACA Lewis 8- by 6-foot super-
sonic tunnel. Comparison of experimental station
normal force with those predicted by linearized po-
tential theory (on which radius of body assumed to
approach zero) shows that inaccurate prediction by
theory of normal force acting on slender body of rev-
olution with curved profiles at angle-of-attack re-
sults from inaccurate prediction of potential flow
pressure distribution due to angle of attack, and
from neglecting effect of cross-flow separation. In-
crease in total axial force with angle of attack was
primarily due to increase in base pressure force.
NACA RM E51H23
WIRE CLOTH AS POROUS MATERIAL FOR
TRANSPIRATION-COOLED WALLS. E. R. G.
Eckert, Martin R. Kinsler, and Reeves P. Cochran.
November 1951. 38p. diagrs., photos., tab. (NACA
RM E51H23) (Declassified from Confidential,
9/15/55)
The permeability characteristics and tensile strength
of a porous material developed from stainless-steel
corduroy wire cloth for use in transpiration-cooled
vvalls where the primary stresses are in one direc-
tion were investigated. The results of this investi-
gation are presented and compared with similar re-
sults obtained with porous sintered metal compacts.
A much wider range of permeabilities is obtainable
with the wire cloth than with the porous metal com-
pacts considered and the ultimate tensile strength in
the direction of the primary stresses for porous ma-
terials produced from three mesh sizes of wire cloth
is from two to three times the ultimate tensile
strengths of the porous metal compacts.
NACA RM E52I12
INTERSTAGE SURVEYS AND ANALYSIS OF VIS-
COUS ACTION IN LATTER STAGES OF A MULTI-
STAGE AXIAL-FLOW COMPRESSOR. William B.
Briggs and Charles C. Giamati. March 1953. 51p.
diagrs., photo., tab. (NACA RM E52I12)
(Declassified from Confidential, 9/15/55)
The overall performance of an eight-stage axial-
flow compressor having a design stage pressure
ratio of 1.23 was determined and radial interstage
surveys were made for four weight flows at both 50
percent and 100 percent design speed. The survey
data, from behind the guide vanes and each row of
NACA
RESEARCH ABSTRACTS NO. 91
the fifth to eighth stages, were presented as axial
and tangential velocity components from which
boundary-layer parameters of displacement and
momentum thickness were calculated. An analysis
related current qualitative ideas of flows which are
imposed upon the through flow and the quantitative
status of theory to the observed variation of these
parameters. Midchannel velocity peaks in excess
of design value were not found.
NACA RM E51I25
FLOW SEPARATION AHEAD OF A BLUNT AXDALLY
SYMMETRIC BODY AT MACH NUMBERS 1.76 TO
2.10. W. E. Moeckel. December 1951. 12p.
diagrs., photos. (NACA RM E51I25) (Declassified
from Confidential, 9/15/55)
The pressure distribution and drag were determined
for a spherical-nosed axially symmetric body with
thin projecting rods at Mach numbers of 1.76, 1.93,
and 2.10. The upstream projection distance of the
rods was varied over a wide range to study changes
in the character of the flow separation and to deter-
mine the variation of drag and pressure distribution
with tip projection. Drag coefficients between 0.18
and 0.30 were obtained for most tip projections at
each Mach number.
NACA RM L8I08
LONGITUDINAL-STABILITY INVESTIGATION OF
HIGH-LIFT AND STALL-CONTROL DEVICES ON A
52° SWEPTBACK WING WITH AND WITHOUT FUSE-
LAGE AND HORIZONTAL TAIL AT A REYNOLDS
NUMBER OF 6.8 x 10 6 . Gerald V. Foster and
James E. Fitzpatrick. December 20, 1948. 41p.
diagrs., photos., tabs. (NACA RM L8I08)
(Declassified from Confidential, 9/15/55)
Contains low-speed longitudinal stability character-
istics of a 52° sweptback wing of aspect ratio 2.88,
taper ratio 0.625, and NACA 64j-112 airfoil sections
normal to the 0.282-chord line, in combination with
split flaps, leading-edge flaps, and upper-surface
fences. Low-wing and midwing-fuselage aerody-
namic characteristics are presented with and without
a horizontal tail at various vertical locations. Tests
were conducted at a Reynolds number of 6.8 x 10".
NACA RM L8I30a
EFFECTS OF A SWEPTBACK HYDROFOIL ON THE
FORCE AND LONGITUDINAL STABILITY CHARAC-
TERISTICS OF A TYPICAL HIGH-SPEED AIR-
PLANE. Raymond B. Wood. December 2, 1948.
19p. diagrs., photo., tabs. (NACA RM L8I30a)
(Declassified from Confidential, 9/15/55)
An investigation was conducted in the Langley 8-foot
high-speed tunnel to determine the effects of a swept-
back hydrofoil on the force and longitudinal stability
NACA
RESEARCH ABSTRACTS NO. 9)
characteristics of a typical high-speed airplane.
The Mach number range for this investigation was
from 0.60 to 0.95 and at M = 1.20. The effects of
the hydrofoil on the lift, drag, and pitching-moment
characteristics are presented.
NACA RM L9I30
FLIGHT INVESTIGATIONS AT HIGH-SUBSONIC,
TRANSONIC, AND SUPERSONIC SPEEDS TO DE-
TERMINE ZERO-LIFT DRAG OF FIN-STABILIZED
BODIES OF REVOLUTION HAVING FINENESS
RATIOS OF 12.5, 8.91, AND 6.04 AND VARYING
POSITIONS OF MAXIMUM DIAMETER. Roger G.
Hart and Ellis R. Katz. November 30, 1949. 36p.
diagrs., photos. (NACA RM L9I30)
(Declassified from Confidential, 9/15/55)
Rocket-powered models were flown at transonic and
supersonic speeds to determine the zero-lift drag of
fin-stabilized bodies of revolution differing in fine-
ness ratio and in position of maximum diameter.
The bodies were of fineness ratio 12.5, 8.91, and
6.04 and all had cut-off sterns with equal base and
frontal areas.
NACA RM L50I27
THE LONGITUDINAL STABILITY, CONTROL EF-
FECTIVENESS, AND CONTROL HINGE-MOMENT
CHARACTERISTICS OBTAINED FROM A FLIGHT
INVESTIGATION OF A CANARD MISSILE CONFIG-
URATION AT TRANSONIC AND SUPERSONIC
Sl'EEDS. Roy J. Niewald and Martin T. Moul.
November 24, 1950. 43p. diagrs., photos.
(NACA RM L50I27)
(Declassified from Confidential, 9/15/55)
A 60° delta-wing canard missile configuration was
flight-tested at the Langley Pilotless Aircraft
Research Station at Wallops Island, Va. Longitudi-
nal stability derivatives, control hinge -moment, and
drag characteristics were obtained at transonic and
supersonic velocities by utilizing a continuous step
control disturbance of ±5°.
NACA RM L51G31
SYSTEMATIC TWO-DIMENSIONAL CASCADE TESTS
OF NACA 65-SERIES COMPRESSOR BLADES AT
LOW SPEEDS. L. Joseph Herrig, James C. Emery,
and John R. Erwin. September 1951. 223p.
diagrs., photo., tabs. (NACA RM L51G31)
(Declassified from Confidential, 9/15/55)
A two-dimensional low-speed porous-wall cascade
tunnel investigation has been conducted to establish
the performance of the NACA 65-series compressor
blade sections over the useful range of inlet angle,
solidity, and section camber. Design points for opti-
mum high-speed operation are presented. The load-
ing limitation is determined for some conditions.
Trends of section operating range with increasing
section camber are determined for the four inlet
angles tested.
NACA RM L51H03
ROLLING EFFECTIVENESS OF ALL-MOVABLE
WINGS AT SMALL ANGLES OF INCIDENCE AT
MACH NUMBERS FROM 0.6 TO 1.6. H. Kurt
Strass and Edward T. Marley. October 1951. 16p.
diagrs., photo., tab. (NACA RM L51H03)
(Declassified from Confidential, 9/15/55)
An investigation of the rolling effectiveness of sev-
eral all-movable wing configurations has been con-
ducted throughout a Mach number range from 0.6 to
1.6 in order to check a simplified wing incidence cor-
rection theory which states that for all-movable wing
pb 2i w (\ + 2x\ , ....
5^ = * - h o ( i oy )• In addition, a comparison was
made with two other more complex methods of esti-
mation. The results showed that the simplified
theory gave accurate agreement with experiment at
Mach numbers from 0.6 to 1.6 and suggest the use of
this simplified equation as a means of predicting the
rolling effectiveness of all-movable wings throughout
this speed range.
NACA RM L52I16
SOME EXPERIMENTAL STUDIES OF PANEL
FLUTTER AT MACH NUMBER 1.3. Maurice A.
Sylvester and John E. Baker. December 1952.
25p. diagrs., photos., tab. (NACA RM L52I16)
(Declassified from Confidential, 9/15/55)
Experimental studies of panel flutter using thin metal
plates .vere conducted at a Mach number of 1.3 to
verify its existence and to study the effects of some
structural parameters on the flutter characteristics.
The effects of tensile forces and buckling were
studied on panels clamped front and rear, in addition
to initially buckled panels clamped on all four edges.
Panel flutter was obtained under these laboratory
conditions and it was found that tensile forces,
shortening the panels, and increasing the bending
stiffness were effective means for eliminating flutter.
Buckled panels were more susceptible to flutter than
unbuckled panels. No apparent systematic trends in
the flutter modes or frequencies could be observed.
NACA RM L53F17
METHOD OF ESTIMATING THE INCOMPRESSIBLE-
FLOW PRESSURE DISTRIBUTION OF COMPRESSOR
BLADE SECTIONS AT DESIGN ANGLE OF ATTACK.
John R. Erwin and Laura A. Yacobi. December
1953. 41p. diagrs., tab. (NACA RM L53F17)
(Declassified from Confidential, 9/15/55)
A method was devised for estimating the incom-
pressible flow pressure distribution over com-
pressor blade sections at design angle of attack.
The theoretical incremental velocities due to camber
and thickness of the section as an isolated airfoil
are assumed proportional to the average passage
velocity and are modified by empirically determined
interference factors. Comparisons were made be-
tween estimated and test pressure distributions of
NACA 65-series sections for typical conditions.
Good agreement was obtained.
10
NACA RM L53I11
UNIVERSITY OF FLORIDA
3 1262 08153 283 9
NACA
RESEARCH
ABSTRACTS NO. 91
FREE-FLIGHT -TUNNEL INVESTIGATION OF THE
LOW-SPEED STABILITY AND CONTROL CHARAC-
TERISTICS OF A CANARD AIRPLANE MODEL.
Joseph L. Johnson, Jr., and John W. Paulson.
October 1953. 37p. diagrs., photo., 2 tabs. (NACA
RM L53I11) (Declassified from Confidential,
9/15/55)
Results are presented of an experimental investiga-
tion in the Langley free-flight tunnel to determine
the dynamic lateral stability and control character-
istics of a model of a canard-type airplane. Tests
were made with several vertical-tail configurations
for the model with a triangular horizontal tail and
with a sweptback horizontal tail having a leading-
edge flap.
NACA RM L53I25a
FLIGHT DETERMINATION OF DRAG OF NORMAL-
SHOCK NOSE INLETS WITH VARIOUS COWLING
PROFILES AT MACH NUMBERS FROM 0.9 TO 1.5.
R. I. Sears, C. F. Merlet, and L. W. Putland.
October 1953. 36p. diagrs., photos., tabs. (NACA
RM L53I25a) (Declassified from Confidential,
9/15/55)
External-drag data are presented for normal-shock
nose inlets with 1 -series, parabolic, and conic
cowling profiles. The tests were made at an angle
of attack of 0° by using rocket -propelled models in
free flight at Mach numbers from 0.9 to 1.5. The
Reynolds number based on body maximum diameter
varied from 2.5 x 10° to 5.5 x 10°\ At maximum
flow rate, the inlet models had about the same ex-
ternal drag at a Mach number of approximately 1.1,
but at higher Mach numbers the sharp-lip conic cowl
had the least drag. Blunting or beveling the lip of
the conic cowl while keeping the fineness ratio
constant resulted in drag coefficients slightly higher
than for the sharp-lip conic cowl at maximum flow
rate. At a mass-flow ratio of about 0.8, the conic
cowls with sharp, blunt, or beveled lips and the
parabolic cowl all gave about the same drag.
NACA RM L53I29b
INVESTIGATION AT SUPERSONIC SPEEDS OF THE
VARIATION WITH REYNOLDS NUMBER AND MACH
NUMBER OF THE TOTAL, BASE, AND SKIN-
FRICTION DRAG OF SEVEN BOATTAIL BODIES OF
REVOLUTION DESIGNED FOR MINIMUM WAVE
DRAG. August F. Bromm, Jr.. and Julia M.
Goodwin. December 1953. 20p. diagrs., photo.
(NACA RM L53I29b) (Declassified from
Confidential, 9/15/55)
Results are presented from an investigation of the
variation with Reynolds number and Mach number of
the total, base, and skin-friction drag of seven boat-
tail bodies of revolution designed for minimum wave
drag according to the theory of NACA TN 2550. The
tests covered a Reynolds number range from approxi-
mately 1.0 x 10 6 to 10.0 x 10 6 at Mach numbers of
1.62, 1.93, and 2.41, respectively.
NACA RM L53I30b
TWO-DIMENSIONAL LOW -SPEED CASCADE INVES-
TIGATION OF NACA COMPRESSOR BLADE SEC-
TIONS HAVING A SYSTEMATIC VARIATION IN
MEAN-LINE LOADING. John R. Erwin, Melvyn
Savage, and James C. Emery. November 1953.
129p. diagrs., tabs. (NACA RM L53I30b)
(Declassified from Confidential, 9/15/55)
The low-speed cascade performance of the high-
speed NACA 65- (C-i A2l8b) 1( ^ compressor blade
sections has been systematically investigated. When
used in conjunction with published cascade data, the
results will provide design information for all inlet
angle and solidity conditions within the usual range of
application. Summary curves have been prepared to
facilitate the selection of blade sections and settings
to fulfill the conditions dictated by compressor design
velocity diagrams. Comparative tests of blade
sections having widely different loading distributions
indicated that these data, in conjunction with previ-
ously published cascade data, permit a fairly
accurate prediction of design performance for most
compressor blade sections since the mean lines
tested probably encompass the practical range of
compressor-blade mean-line loading distributions.
A comparative evaluation of the high-speed per-
formance capabilities of the blade sections investi-
gated was made.
NACA - Langley Field, Va.
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